Inner shroud cooling arrangement in a gas turbine engine

ABSTRACT

A component in a gas turbine engine includes an airfoil and a shroud. The shroud has an outer surface supporting an end of the airfoil and defines a portion of an annular gas path. The shroud includes axial edges extending between upstream and downstream edges thereof. Each of the axial edges includes a seal slot that receives a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends between the upstream and downstream edges of the shroud. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot.

FIELD OF THE INVENTION

The present invention relates to turbine engines and, more particularly, to cooling arrangements for inner shrouds of vane segments in gas turbine engines.

BACKGROUND OF THE INVENTION

In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor section then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., turbine blades, where the turbine blades extract energy from the hot combustion gases for powering the compressor section and providing output power. Because the airfoils, i.e., vanes and turbine blades, are directly exposed to the hot combustion gases, they are typically provided with internal cooling channels that may feed a cooling fluid, such as compressor bleed air, through the airfoil and through various passages formed in structure associated with the vanes and/or blades.

One type of stationary airfoil in a turbine engine is provided as a component of a stator vane segment. The stator vane segment may include a radially inner shroud, a radially outer shroud, and one or more airfoils extending between the inner and outer shrouds. Hot combustion gases, or working gases, may be supplied from a combustor section and pass through passages defined between adjacent airfoils and between the inner and outer shrouds, resulting in some of the heat of the gases being transferred to the vane segments. As turbine engine performance has been increased with increasing combustion gas temperature, there has been a continuing need to improve cooling to the various portions of vane segments in order to avoid or minimize deterioration of the material forming the vane segments.

SUMMARY OF THE INVENTION

In accordance with a first aspect of the present invention, a component is provided in a gas turbine engine. The component comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream edge and the downstream edge. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream edge and the downstream edge. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to the one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.

In accordance with a second aspect of the present invention, a vane is provided in a gas turbine engine. The vane comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream edge and the downstream edge. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream edge and the downstream edge. A cooling air supply passage extends from a cooling air chamber at inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to the one of the axial edges, the cooling air exit passage comprising a purge passage having an exit opening providing a volume of air for purging hot gas from the seal slot and the seal. The exit opening is located at a downstream end of the cooling air channel adjacent the downstream edge of the shroud. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.

In accordance with a third aspect of the present invention, a vane is provided in a gas turbine engine. The vane comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream and downstream edges. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream and the downstream edges. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. A plurality of impingement passages extend from the cooling air channel to the one of the axial edges, the impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of the adjacent shroud, the exit openings being located at an upstream end of the cooling air channel and adjacent the upstream edge of the shroud. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.

BRIEF DESCRIPTION OF THE DRAWINGS

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

FIG. 1 is a perspective view of a portion of a row of vane segments including inner shroud cooling arrangements according to an embodiment of the invention;

FIG. 2 is a perspective view of a radially inner side of one of the vane segments of FIG. 1 and including cut-away portions illustrating an inner shroud cooling arrangement;

FIG. 3 is a diagrammatic top perspective view looking in a radially inward direction of the vane segment shown in FIG. 2, wherein the outer shroud and airfoil of the vane segment have been removed for clarity;

FIG. 4 is an end view looking in an axial direction illustrating mating edges of two vane segments of the row of vane segments illustrated in FIG. 1;

FIG. 5 is a top view looking in a radially inward direction illustrating mating edges of two vane segments of the row of vane segments illustrated in FIG. 1;

FIG. 6 is a diagrammatic top perspective view looking in a radially inward direction of a vane segment including an inner shroud cooling arrangement in a row of vane segments according to another embodiment of the invention, wherein the outer shroud and airfoils of the vane segment have been removed for clarity;

FIG. 7 is an end view looking in an axial direction illustrating mating edges of the vane segment illustrated in FIG. 6 and an adjacent vane segment; and

FIG. 8 is a top view looking in a radially inward direction illustrating mating edges of the vane segment illustrated in FIG. 6 and an adjacent vane segment.

DETAILED DESCRIPTION OF THE INVENTION

In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

Referring to FIG. 1, a portion of a row 10 of vanes or stator vane segments 12 is illustrated, such as may be incorporated in a turbine section 14 of a turbine engine. The row 10 of vane segments 12 shown in FIG. 1 comprises a first row 10 of vane segments 12, also referred to as row one vane segments, of the turbine section 14. The vane segments 12 each include at least one airfoil 16, a radially inner shroud 18 rigidly connected to a radially inner end of the airfoil 16, and a radially outer shroud 20 rigidly connected to a radially outer end of the airfoil 16. Each airfoil 16 comprises a generally concave pressure side 22 and a generally convex suction side 24. The pressure and suction sides 22, 24 are connected at an upstream leading edge 26 and at a chordally spaced downstream trailing edge 28 (FIG. 3).

Referring to FIG. 1, the outer shroud 20 is suspended radially inwardly from a casing structure (not shown) of the turbine section, and the row 10 of vane segments 12 comprises a plurality of vane segments 12 suspended in side-by-side relation extending circumferentially about a turbine rotor (not shown) within the turbine engine. Although the vane segments 12 are illustrated as each including one airfoil 16, it should be understood that the vane segments 12 may be constructed with two or more airfoils 16. A hot working gas H_(G) created in a conventional combustor assembly (not shown) is discharged into the turbine section 14 and passes between adjacent airfoils 16 of the vane segments 12. The vane segments 12 direct the hot working gas H_(G) toward rows of blades (not shown) in the turbine section, which blades are caused to rotate and cause corresponding rotation of the turbine rotor.

As noted above, the vane segments 12 are suspended in a circumferential row 10 about the turbine rotor, such that the airfoils 16 are spaced apart and define flow passages 30 therebetween for channeling the hot working gas H_(G) through the turbine section 14 during engine operation. Each flow passage 30 forms a portion of an annular path for the hot working gas H_(G) and is bounded by the pressure sidewall 22 of one airfoil 16 and the suction sidewall 24 of an adjacent airfoil 16. The flow passage 30 is also defined between the inner shroud 18 and the outer shroud 20 and extends in a flow direction from an upstream edge 32 to a downstream edge 34 of the inner shroud 18 (FIGS. 2 and 3) and from an upstream edge 36 to a downstream edge 38 of the outer shroud 20 (FIGS. 1 and 2).

Each vane segment 12 includes a first generally axially extending mating edge 40 extending between the upstream edge 32 and the downstream edge 34 of the inner shroud 18, and an opposing second generally axially extending mating edge 42 extending generally parallel to the first mating edge 40 between the upstream edge 32 and the downstream edge 34 of the inner shroud 18. Each of the mating edges 40, 42 includes a generally axially extending seal slot 44 adapted to receive an axially extending seal member 46 (FIG. 1). The seal member 46 spans across a gap formed between adjacent mating edges 40, 42 of adjacent vane segments 12 to separate the hot working gas H_(G) from a region of cool air located radially inwardly from the inner shroud 18.

Referring to FIGS. 2-5, respective primary cooling air channels 48 extend within the inner shroud 18 generally axially adjacent to and substantially parallel to each of the mating edges 40, 42. As illustrated in FIG. 3, in the embodiment shown the channels 48 extend from upstream portions 48A thereof adjacent to the upstream edge 32 of the inner shroud 18 to outlets 49 located at downstream portions 48B thereof adjacent to the downstream edge 34 of the inner shroud 18. Each channel 48 is located radially between an outer surface 50 defined by an outer wall 51 of the inner shroud 18 and the seal slot 44, see FIG. 4. Further, each channel 48 is preferably aligned circumferentially with a respective seal slot 44, such that the channels 48 are located on, or intersected by, respective radial planes 54 passing radially through the corresponding seal slots 44, see FIG. 4. Preferably, the entirety of a cross-sectional area of the cooling air channel 48, as viewed in the axial direction, is located radially outwardly from, i.e., substantially overlies, a respective seal slot 44, as shown in FIG. 4. The cooling air channels 48 effect convective cooling of a corner 56 at an intersection of the outer surface 50 of the outer wall 51 and one of the mating edges 40, 42 of the inner shroud 18, as will be described herein. It is noted that the spacing between the outer surface 50 of the outer wall 51 and the seal slots 44 is large enough to accommodate the respective cooling air channels 48 therebetween while still maintaining the structural rigidity of the inner shroud 18 at the respective mating edges 40, 42.

Referring now to FIGS. 2 and 3, an inner surface 52 (FIGS. 2 and 4) of the outer wall 51 defines a radially outer boundary for a leading edge cooling air chamber 53 associated with the inner shroud 18. The chamber 53 receives cooling air, i.e., compressor discharge air, for cooling the inner shroud 18 proximate to the upstream edge 32 thereof. In the embodiment shown, cooling air enters the chamber 53 from a first inner cavity 55 (FIGS. 1 and 2) located radially inwardly from the inner shroud 18 proximate to the upstream edge 32 thereof. The cooling air enters the chamber 53 through impingement holes 57A formed in an impingement plate 57, which impingement plate 57 defines a radially inner boundary for the chamber 53, see FIG. 2. The cooling air passing through the holes 57A impinges on the inner surface 52 of the outer wall 51, and then a portion of the cooling air flows through film cooling holes 59 (FIGS. 1 and 3) formed in the outer wall 51. The cooling air passing through the film cooling holes 59 provides film cooling for the outer surface 50 of the outer wall 51 and may then be swept up and carried through the annular path with the hot working gas H_(G).

As shown in FIGS. 2 and 3, primary cooling air supply passages 58 provide fluid communication between the leading edge cooling air chamber 53 and the upstream portions 48A of the respective cooling air channels 48 for supplying a first portion of cooling air from the chamber 53 to the channels 48. The passages 58 may include a component in the radial direction, since the channels 48 in the embodiment shown may be located in the outer wall 51 of the inner shroud 18 radially outwardly from the chamber 53. Further, the passages 58 may be angled toward the downstream edge 34 of the inner shroud 18 to promote a flow toward the downstream portions 48B of the cooling air channels 48.

The inner surface 52 of the outer wall 51 further defines a radially outer boundary for a mid-chord cooling air chamber 60 (FIGS. 2-4).

The mid-chord cooling air chamber 60 receives cooling air, i.e., compressor discharge air, for cooling a mid-chord portion 18A and a trailing edge portion 18B of the inner shroud 18, see FIGS. 2 and 3. In the embodiment shown, cooling air enters the chamber 60 from a second inner cavity 62 (FIG. 2) located radially inwardly from the inner shroud 18 proximate to the mid-chord portion 18A thereof. The cooling air enters the chamber 60 through impingement holes 63A formed in a plurality of impingement plates 63 associated with the mid-chord portion 18A, which impingement plates 63 define a radially inner boundary for the mid-chord cooling air chamber 60, see FIG. 2. It is noted that additional or fewer impingement plates 63 may be used than as shown in FIG. 2 to define additional or fewer sub-chambers of the mid-chord cooling air chamber 60. For example, the mid-chord cooling air chamber 60 illustrated diagrammatically in FIG. 3 has been consolidated into a single chamber, whereas the mid-chord cooling air chamber 60 illustrated in FIG. 2 comprises a plurality of sub-chambers that collectively form the mid-chord cooling air chamber 60. A portion of the cooling air passing through the holes 63A impinges on the inner surface 52 of the outer wall 51, and then flows through film cooling holes 64 (FIGS. 1 and 3) formed in the outer wall 51. The cooling air passing through the film cooling holes 64 provides film cooling for the outer surface 50 of the outer wall 51 and may then be swept up and carried through the annular path with the hot working gas H_(G).

The portion of the mid-chord cooling air chamber 60 associated with the trailing edge portion 18B of the inner shroud 18 is associated with a cover plate 66 (FIG. 2) that covers and defines a radially inner boundary for the trailing edge portion 18B of the mid-chord cooling air chamber 60. The cooling air passing into the trailing edge portion 18B of the mid-chord cooling air chamber 60 passes over a bridge 65 that spans between the mating edges 40, 42, i.e., between the outer wall 51 and the bridge 65. The cooling air in the trailing edge portion 18B of the mid-chord cooling air chamber 60 provides convective cooling for the trailing edge portion 18B of the inner shroud 18, and then flows through film cooling holes 67 (FIG. 3) formed in the outer wall 51. The cooling air passing through the film cooling holes 67 provides film cooling for the outer surface 50 of the outer wall 51 at the trailing edge portion 18B of the inner shroud 18 and may then be swept up and carried through the annular path with the hot working gas H_(G).

Referring to FIGS. 2-5, replenishing cooling air supply passages 68 provide fluid communication between the mid-chord cooling air chamber 60 and the respective cooling air channels 48 at a location between the upstream and downstream portions 48A and 48B thereof, i.e., between the primary cooling air supply passages 58 and the downstream edge 34 of the inner shroud 18, for supplying a second portion of cooling air from the mid-chord cooling air chamber 60 to the channels 48. As shown in FIG. 4, the passages 68 may include a component in the radial direction, since the channels 48 in the embodiment shown may be located in the outer wall 51 of the inner shroud 18 radially outwardly from the mid-chord cooling air chamber 60. Further, the passages 68 may be angled toward the downstream edge 34 of the inner shroud 18 to promote a flow toward the downstream portions 48B of the cooling air channels 48, as shown in FIG. 5.

As shown in FIGS. 2-5, one or more cooling air exit passages 69 extend from each cooling air channel 48 to the respective axial mating edge 40, 42. The cooling air exit passages 69 according to this embodiment comprise purge passages having a diameter D₁ (see FIG. 4) that is large enough, e.g., 3-4 mm, for providing a sufficient volume of cooling air for purging hot gas from the seal slot 44 and the seal member 46 and for creating a barrier of cool air in the gap between adjacent inner shrouds 18.

Each cooling air exit passage 69 is in fluid communication with the cooling air channel 48 and includes an exit opening 70 located between the outer surface 50 of the inner shroud 18 and the seal slot 44 (see FIGS. 2 and 4). As shown in FIGS. 2 and 3, the cooling air exit passages 69 according to this embodiment are located along a downstream region 71 of the mating edges 40, 42, and are preferably located toward the downstream portions 48B of the cooling air channels 48, i.e., closer to the downstream edge 34 of the inner shroud 18 than to the upstream edge 32 thereof, although the cooling air exit passages 69 could be located elsewhere along the respective mating edges 40, 42 as desired. Further, at least one of the cooling air exit passages 69 associated with each cooling air channel 48 is located near a respective replenishing cooling air supply passage 68. As illustrated in FIG. 5, the passages 69 may be angled toward the downstream edge 34 of the inner shroud 18 to promote a flow having a velocity component in the same direction as the hot working gas H_(G). Moreover, the exit openings 70 of the passages 69 may be axially offset with respect to exit openings 70 of adjacent inner shrouds 18, as shown in FIG. 5.

During operation of the engine, the row 10 of vane segments 12, which, as noted above, is a first row 10 of vane segments 12 according to this embodiment, is exposed to the high temperature hot working gas H_(G) entering the turbine section 14 from the combustor assembly. A region of the vane segment 12 along the inner shroud 18 adjacent to and downstream from the trailing edge 28 is believed to be particularly susceptible to exposure to high temperature gases that may lead to elevated temperatures of the surfaces of the inner shroud 18. In particular, the mating edges 40, 42 may be exposed to elevated temperatures in the region 71, which may generally be defined as extending from a location at or near an axial location of the trailing edge 28, as identified by line 73 in FIG. 3, to the location of the downstream edge 34. The replenishing cooling air supply passages 69, air exit passages 69 and exit openings 70 are preferably located in or near the region 71.

Cooling air, e.g., compressor discharge air, enters the leading edge cooling air chamber 53 from the first cavity 55 through the impingement holes 57A formed in the impingement plate 57. The cooling air entering the chamber 53 through the impingement holes 57A contacts and provides impingement cooling to the inner surface 52 of the outer wall 51.

A first portion of the cooling air in the leading edge cooling air chamber 53 passes into the cooling air channels 48 through the respective primary cooling air supply passages 58. Another portion of the cooling air in the chamber 53 passes through the film cooling holes 59 in the wall 51 and provides film cooling for the outer surface 50 of the wall 51 as discussed above.

The cooling air flowing through the cooling air channels 48 provides convective cooling for the inner shroud 18 adjacent to the mating edges 40, 42. Since the cooling air channels 48 are located close to the mating edges 40, 42 and to the outer surface 50 of the wall 51, the corners 56 of the inner shroud 18 are convectively cooled by the cooling air flowing through the cooling air channels 48.

As a result of the cooling air in the cooling air channels 48 convectively cooling the inner shroud 18 adjacent to the mating edges 40, 42 by removing heat from the inner shroud 18, the cooling air flowing through the cooling air channels 48 is heated as it flows downstream toward the downstream portions 48B of the cooling air channels 48.

The second portion of cooling air, which is provided into the cooling air channels 48 from the mid-chord cooling air chamber 60 via the replenishing cooling air supply passages 68, is added to and cools the first portion of cooling air. Hence, the cooling air that is available for convective cooling in the downstream portions 48B of the cooling air channels 48, i.e., for cooling the mating edges 40, 42 and the corners 56, is cool enough to sufficiently cool the inner shroud 18 adjacent to the downstream edge 34 thereof. That is, the cooling air exiting the cooling air exit passages 69 through the exit openings 70 is cool enough, and comprises an adequate volume, to sufficiently cool the seal member 46 and the mating edges 40, 42 and corners 56 of the inner shroud 18, and to provide a cool air barrier for the hot working gases H_(G). Further, since the exit openings 70 of the respective adjacent inner shrouds 18 are axially offset from one another, the cooling air exiting the respective exit openings 70 provides a substantially even distribution of cooling air between the mating edges 40, 42 of the respective inner shrouds 18. It is noted that any cooling air in the channels 48 that does not exit through the cooling air exit passages 69 may exit the inner shroud 18 through the outlets 49 of the channels 48.

Referring now to FIGS. 6-8, a vane segment 110 belonging to a row (not shown in this embodiment) of vane segments 110 according to another embodiment of the invention is shown. The vane segment 110 according to this embodiment belongs to a second row of vane segments 110, also referred to as a row two vane segment. The vane segment 110 illustrated in FIG. 6 includes two airfoils 112 and a radially inner shroud 114 rigidly connected to a radially inner end of each airfoil 112. The vane segment 110 also includes one or more radially outer shrouds (not shown in this embodiment) rigidly connected to radially outer ends of the airfoils 112. It is noted that the vane segment 110 may be constructed with additional or fewer airfoils 112.

The vane segment 110 includes a first generally axially extending mating edge 116 extending between an upstream edge 118 and a downstream edge 120 of the inner shroud 114, and an opposing second generally axially extending mating edge 122 extending generally parallel to the first mating edge 116 between the upstream edge 118 and the downstream edge 120 of the inner shroud 114. Each of the mating edges 116, 122 includes a generally axially extending seal slot 124 adapted to receive an axially extending seal member 126 (FIG. 7). The seal member 126 spans across a gap formed between adjacent mating edges 116, 122 of adjacent vane segments 110 to separate the hot working gas (discussed above with reference to FIGS. 1-5) from a region of cool air located radially inwardly from the inner shroud 114.

As shown in FIG. 6-8, respective primary cooling air channels 130 extend within the inner shroud 114 generally axially adjacent to and substantially parallel to each of the mating edges 116, 122. As illustrated in FIG. 6, in the embodiment shown the channels 130 extend from upstream portions 130A thereof adjacent to the upstream edge 118 of the inner shroud 114 to outlets 131 located at downstream portions 1308 thereof adjacent to the downstream edge 120 of the inner shroud 114. As shown in FIG. 7, each channel 130 is located radially between an outer surface 132 defined by an outer wall 134 of the inner shroud 114 and the seal slot 124. Further, each channel 130 is preferably aligned circumferentially with a respective seal slot 124, such that the channels 130 are located on, or intersected by, respective radial planes 136 passing radially through the corresponding seal slots 124, see FIG. 7. Preferably, the entirety of a cross-sectional area of the cooling air channel 130, as viewed in the axial direction, is located radially outwardly from, i.e., substantially overlies, a respective seal slot 124, as shown in FIG. 7. The cooling air channels 130 effect convective cooling of a corner 138 (FIG. 7) at an intersection of the outer surface 132 of the outer wall 134 and respective ones of the mating edges 116, 122 of the inner shroud 114, as will be described herein. It is noted that the spacing between the outer surface 132 of the outer wall 134 and the seal slots 124 is large enough to accommodate the respective cooling air channels 130 therebetween while still maintaining the structural rigidity of the inner shroud 114 at the respective mating edges 116, 122.

Referring now to FIGS. 6 and 7, an inner surface 140 (FIG. 7) of the outer wall 134 defines a radially outer boundary for a mid-chord cooling air chamber 142 associated with the inner shroud 114. The chamber 142 receives cooling air, i.e., compressor discharge air, for cooling the inner shroud 114. In the embodiment shown, cooling air, i.e., compressor discharge air, may enter the chamber 142 from internal cooling passageways 144 (FIG. 6) that extend through the airfoils 112. It is noted that, while film cooling holes (such as the film cooling holes 64 described above with reference to FIGS. 1 and 3) are not illustrated in the outer wall 134 of the inner shroud 114 according to this embodiment, such film cooling holes could be provided in the outer wall 134 if film cooling of the outer surface 132 of the outer wall 134 is desired.

As shown in FIG. 6, cooling air is provided from the mid-chord cooling air chamber 142 via passageways 148 into a leading edge cooling air chamber 150 located adjacent to the upstream edge 118 of the inner shroud 114. The cooling air in the leading edge cooling air chamber 150 provides convective cooling to the inner shroud 114 adjacent to the upstream edge 118 thereof. It is noted that the cooling air chamber 150 may be formed by one or more sub-chambers, two such sub-chambers define the leading edge cooling air chamber 150 in the embodiment shown in FIG. 6.

A first portion of cooling air is provided into the cooling air channels 130 via respective primary cooling air supply passages 152 that extend from the leading edge cooling air chamber 150 to the upstream portions 130A of the respective cooling air channels 130, see FIG. 6. The passages 152 may include a component in the radial direction, as the channels 130 in the embodiment shown may be located in the outer wall 134 of the inner shroud 114 radially outwardly from the leading edge cooling air chamber 150.

Referring to FIGS. 6-8, replenishing cooling air supply passages 156 provide fluid communication between the mid-chord cooling air chamber 142 and the respective cooling air channels 130 between the upstream and downstream portions 130A and 130B thereof, i.e., between the primary cooling air supply passages 152 and the downstream edge 120 of the inner shroud 114. The replenishing cooling air supply passages 156 supply a second portion of cooling air from the mid-chord cooling air chamber 142 to the channels 130. As shown in FIG. 7, the passages 156 may include a component in the radial direction, as the channels 130 in the embodiment shown may be located in the outer wall 134 of the inner shroud 114 radially outwardly from the chamber 142. Further, the passages 156 may be angled toward the downstream edge 120 of the inner shroud 114 to promote a flow toward the downstream portions 130B of the cooling air channels 130, as shown in FIGS. 6 and 8.

As shown in FIGS. 6-8, a plurality of cooling air exit passages 158 extend from each cooling air channel 130 to the respective axial mating edges 116, 122. The cooling air exit passages 158 according to this embodiment comprise impingement passages having a diameter D₂ (see FIG. 7) that is large enough, e.g., 1-2 mm, for providing impingement cooling to a mating edge 116, 122 of an adjacent vane segment 110, see FIGS. 7 and 8.

Each cooling air exit passage 158 is in fluid communication with a respective cooling air channel 130 and includes an exit opening 160 located between the outer surface 132 of the inner shroud 114 and the seal slot 124 (see FIG. 7). As shown in FIG. 6, the cooling air exit passages 158 according to this embodiment are preferably located near the upstream portions 130A of the channels 130, i.e., from the upstream edge 118 of the inner shroud 114 to a location about ⅓ of the distance between the upstream and downstream edges 118 and 120 of the inner shroud 114, although it is understood that the cooling air exit passages 158 could be located elsewhere at the mating edges 116, 122 as desired. As illustrated in FIGS. 6 and 8, the passages 158 may be angled toward the downstream edge 120 of the inner shroud 114 to promote a flow including a velocity component in the same direction as the hot working gas. Moreover, the exit openings 160 of the passages 158 may be axially offset with respect to the exit openings 160 of passages 158 of the adjacent inner shroud 114, as shown in FIG. 8.

During operation of the engine, the row of vane segments 110, which, as noted above, is a second row of vane segments 110 according to this embodiment, is exposed to high temperature hot working gas entering the turbine section from the combustor assembly as described above with reference to FIGS. 1-5. A region of the vane segment 110 along the inner shroud 114 adjacent to and downstream from the upstream edge 118, i.e., in a region extending approximately ⅓ of the distance between the upstream and downstream edges 118 and 120, is particularly susceptible to high temperatures.

Cooling air, e.g., compressor discharge air, enters the mid-chord cooling air chamber 142 from the internal cooling passageways 144 that extend through the airfoils 112. The cooling air entering the mid-chord cooling air chamber 142 provides convective cooling to the inner shroud 114 around the chamber 142.

As noted above, cooling air is provided from the mid-chord chamber cooling air chamber 142 into the leading edge cooling air chamber 150 through the passageways 148. The cooling air in the leading edge cooling air chamber 150 provides convective cooling to the inner shroud 114 adjacent to the upstream edge 118 thereof. A first portion of the cooling air in the leading edge cooling air chamber 150 passes into the cooling air channels 130 through the respective primary cooling air supply passages 152.

The cooling air flowing through the cooling air channels 130 provides convective cooling for the inner shroud 114 adjacent to the mating edges 116, 122. Since the cooling air channels 130 are located close to the mating edges 116, 122 and to the outer surface 132 of the wall 134, the corners 138 of the inner shroud 114 are convectively cooled by the cooling air flowing through the cooling air channels 130. As a result of the cooling air in the channels 130 convectively cooling the inner shroud 114 adjacent to the mating edges 116, 122, the cooling air flowing through the cooling air channels 130 is heated as it flows downstream toward the downstream portions 130B of the cooling air channels 130. Some of the first portion of cooling air exits the channels 130 through the cooling air exit passages 158 near the upstream portions 130A of the channels 130.

The second portion of cooling air, which is provided to the cooling air channels 130 directly from the mid-chord cooling air chamber 142 via the replenishing cooling air supply passages 156, is added to and cools the remaining portion of the first portion of cooling air. Hence, the cooling air that is available for cooling within the cooling air channels 130 is cool enough to sufficiently cool the inner shroud 114, i.e., the corners 138 and the mating edges 116, 122, and also to cool the seal members 126. Additionally, since the exit openings 160 of the adjacent inner shrouds 114 are axially offset from one another, the cooling air exiting the respective exit openings 160 provides a substantially even distribution of cooling air between the mating edges 116, 122 of the respective inner shrouds 114 for providing impingement cooling for the opposed mating edges 116, 122. It is noted that any cooling air in the channels 130 that does not exit through the cooling air exit passages 158 may exit the inner shroud 114 through the outlets 131 of the channels 130.

While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention. 

1. A component in a gas turbine engine, said component comprising: an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge; a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge; each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud; a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge; a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel; at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges; and said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
 2. The component of claim 1, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.
 3. The component of claim 1, wherein said cooling air exit passage includes an exit opening located between said outer surface of said shroud and said seal slot.
 4. The component of claim 1, wherein said cooling air exit passage comprises a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud.
 5. The component of claim 4, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.
 6. The component of claim 5, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.
 7. The component of claim 5, wherein said component comprises a vane in a first row of vanes within said gas turbine engine.
 8. The component of claim 1, wherein said cooling air exit passage comprises a plurality of impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel adjacent said upstream edge of said shroud.
 9. The component of claim 8, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.
 10. The component of claim 9, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.
 11. The component of claim 9, wherein said component comprises a vane in a second row of vanes within said gas turbine engine.
 12. A vane in a gas turbine engine, said vane comprising: an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge; a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge; each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud; a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge; a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel; at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges, said cooling air exit passage comprising a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud; and said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
 13. The vane of claim 12, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.
 14. The vane of claim 13, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.
 15. The vane of claim 12, wherein said cooling air channel is without openings for discharge of air along a length of said cooling air channel from said cooling air supply passage to within close proximity of said replenishing cooling air supply passage.
 16. The vane of claim 12, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.
 17. A vane in a gas turbine engine, said vane comprising: an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge; a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream and downstream edges; each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud; a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream and said downstream edges; a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel; a plurality of impingement passages extending from said cooling air channel to said one of said axial edges, said impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel and adjacent said upstream edge of said shroud; and said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
 18. The vane of claim 17, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.
 19. The vane of claim 18, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.
 20. The vane of claim 17, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges. 